Super-cooled ice impact protection for a gas turbine engine

ABSTRACT

A gas turbine engine comprises a fan mounted to rotate about a main longitudinal axis; an engine core, comprising in axial flow series a compressor, a combustor, and a turbine coupled to the compressor through a shaft; a reduction gearbox that receives an input from the shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the shaft; wherein the compressor comprises a first stage at an inlet and a second stage, downstream of the first stage, comprising respectively a first rotor with a row of first blades and a second rotor with a row of second blades, the first and second blades comprising respective leading edges, trailing edges and tips, and wherein the ratio of a maximum leading edge radius of the first blades to a maximum leading edge radius of the second blades is greater than 2.8.

CROSS-REFERENCE TO RELATED APPLICATIONS

This is a Continuation of application Ser. No. 16/686,274 filed Nov. 18,2019, which claims priority to British Application No. 1820423.0 filedon Dec. 14, 2018. The disclosure of the prior application is herebyincorporated by reference herein in its entirety.

BACKGROUND 1. Field of the Disclosure

The present disclosure relates generally to gas turbine engines, andmore specifically to arrangements for super-cooled ice protection for agas turbine engine.

2. Description of the Related Art

Gas turbine engines are used to power aircraft, watercraft, powergenerators, and the like. A gas turbine engine generally comprises, inaxial flow series from front to aft, an air intake, a fan, one or morecompressors, a combustor, one or more turbines, and an exhaust nozzle.Air entering the air intake is accelerated by the fan to produce two airflows: a first air flow (core engine flow) into compressor and a secondair flow (bypass flow) which passes through a bypass duct to providepropulsive thrust. Air entering the compressor is compressed, mixed withfuel and then fed into the combustor, where combustion of the air/fuelmixture occurs. The high temperature and high energy exhaust fluids arethen fed to the turbine, where the energy of the fluids is converted tomechanical energy to drive the compressor in rotation by suitableinterconnecting shaft.

The compressor may be a multi-stage compressor, wherein each compressorstage comprises in axial flow series a row of rotor blades and a row ofstator vanes. A radially inner end of the rotor blades are connected toa hub that define an inner annulus. A casing circumscribes the rotorblades and the stator vanes and defines an outer annulus. The rotorblades and the stator vanes each have a root and an aerofoil portionwith a tip, a trailing edge and a leading edge.

When operating in ice forming conditions (either super-cooled water iceor high altitude ice crystals), ice can accrete on vanes arranged in thecore inlet upstream of the front of the compressor; typically ice isaccreted near the outer annulus. When ice is shed from the vanes, whichmay be owing to aerodynamic loading or vibration, the ice is ingested bythe compressor rotor blade stage immediately downstream.

The severity and size of the ice accretion can be significantly higherfor geared turbofan architectures owing to the low speed and low hubstagger of the fan.

Different anti-icing systems have been proposed to protect the engineagainst ice accretion.

According to one of these approaches, the vanes may be provided withelectrical heaters to prevent ice build-up, and to melt any ice thataccumulates.

Alternatively, relatively hot air bled from the compressor may bedirected towards the vanes.

Both systems are complicated to implement and detrimental to efficiency.

There is therefore a need for an improved system for super-cooled iceimpact protection for a gas turbine engine.

SUMMARY

According to a first aspect, there is provided a gas turbine enginecomprising:

-   -   a fan mounted to rotate about a main longitudinal axis,    -   an engine core, comprising in axial flow series a compressor, a        combustor, and a turbine coupled to the compressor through a        shaft,    -   a reduction gearbox that receives an input from the shaft and        outputs drive to the fan so as drive the fan at a lower        rotational speed than the shaft,        wherein the compressor comprises a first stage at an inlet and a        second stage, downstream of the first stage, comprising        respectively a first rotor with a row of first blades and a        second rotor with a row of second blades, the first and second        blades comprising respective leading edges, trailing edges and        tips,        wherein the ratio of a maximum leading edge radius of the first        blades to a maximum leading edge radius of the second blades is        greater than 2.8.

The blades may comprise an aerofoil portion with a tip, and a root, anda spanwise direction is a direction extending between the tip and theroot of the blades, and a chordwise direction is a direction extendingbetween the leading edge and the trailing edge of the blades.

In the present disclosure, upstream and downstream are with respect tothe air flow through the compressor; and front and rear is with respectto the gas turbine engine, i.e. the fan being in the front and theturbine being in the rear of the engine.

The maximum leading edge radius is a measure of the blade capacity towithstand super-cooled ice impacts: the greater the maximum leading edgeradius, the better the blade capacity to withstand super-cooled iceimpacts.

The present inventors have understood that the first rotor may act as ashield against super-cooled ice impact protecting the downstream stagesof the compressor. To this purpose, the first blades may be maderelatively thicker at the leading edge than the leading edge of thesecond blades to resist super-cooled ice impacts. The advantage ofhaving the first rotor acting as a shield is that the downstream rotorsmay be designed to optimise aerodynamic efficiency, without compromisesand/or penalties linked to super-cooled ice protection. In other words,the downstream rotors, in particular the second rotor, may not need tobe made relatively thicker to resist super-cooled ice impacts, becauseany super-cooled ice entering the compressor may be effectively dealtwith by the first rotor.

Electrical heaters, heat exchangers, or conduits to direct hot airtowards the compressor blades and vanes may therefore be reduced,simplified, or even omitted.

According to the disclosure, the ratio of the maximum leading edgeradius of the first blades the maximum leading edge radius of the secondblades may be greater than 3, or greater than 4, or greater than 5, orgreater than 6, or greater than 7, or greater than 8.

The ratio of the maximum leading edge radius of the first blades to themaximum leading edge radius of the second blades may be less than 10.

For example, the ratio of the maximum leading edge radius of the firstblades to the maximum leading edge radius of the second blades may beless than 9, or less than 8.

The maximum leading edge radius of the first blade may be located in anarea between 70% and 100% of the span height, preferably between 80% and100%, where 0% corresponds to the root and 100% corresponds to the tip.

The maximum leading edge radius of the first blades may be greater than0.4 mm, for example greater than 0.45 mm, or greater than 0.5 mm, orgreater than 0.55 mm, or greater than 0.6 mm.

The maximum leading edge radius of the first blades may be less than 0.9mm, for example less than 0.85 mm, or less than 0.80 mm, or less than0.75 mm, or less than 0.70 mm.

The maximum leading edge radius of the second blades may be comprisedbetween 0.1 and 0.3 mm. For example, the maximum leading edge radius ofthe second blades may be comprised between 0.1 and 0.25 mm, or between0.1 and 0.20 mm, or between 0.15 and 0.30 mm, or between 0.15 and 0.25mm, or between 0.15 mm and 0.20 mm.

The ratio of the maximum leading edge radius of the first blades to aminimum leading edge radius of the first blades may greater than 2.2.For example, the ratio of the maximum leading edge radius of the firstblades to the minimum leading edge radius of the first blades maygreater than 2.5, or greater than 3, or greater than 4, or greater than5.

The ratio of the maximum leading edge radius of the first blades to theminimum leading edge radius of the first blades may be less than 7, forexample less than 6, or less than 5, or less than 4. The ratio of themaximum leading edge radius of the first blades to the minimum leadingedge radius of the first blades may be comprised between 2.2 and 7, forexample between 2.5 and 5, or between 2.5 and 4.

The minimum leading edge radius of the first blade may be located in anarea less than 50% of the span height, for example less than 40%, orless than 30%, or between 15% and 40% of the span height, or between 20%and 30%.

The minimum leading edge radius of the first blades may be greater than0.15 mm, for example greater than 0.20 mm, or greater than 0.21 mm, orgreater than 0.22 mm, or greater than 0.23 mm, or greater than 0.24 mm.

The minimum leading edge radius of the first blades may be less than 0.6mm, for example less than 0.55 mm, or less than 0.5, or less than 0.4mm, or less than 0.35 mm, or less than 0.30 mm.

The ratio of a tip maximum thickness of the second blades to a tipmaximum thickness of the first blades may be less than 0.45. Forexample, the ratio of the tip maximum thickness of the second blades tothe tip maximum thickness of the first blades may be less than 0.40, orless than 0.35, or less than 0.30. For example, the ratio of the tipmaximum thickness of the second blades to the tip maximum thickness ofthe first blades may be greater than 0.20, or greater than 0.25, orgreater than 0.30. For example, the ratio of the tip maximum thicknessof the second blades to the tip maximum thickness of the first bladesmay be comprised between 0.20 and 0.45, or between 0.25 and 0.45, orbetween 0.30 and 0.45, or between 0.35 and 0.45, or between 0.40 and0.45, or between 0.20 and 0.40, or between 0.25 and 0.40, or between0.30 and 0.40.

As the tip is the area of the blades where super-cooled ice impact isgenerally more dangerous to blade integrity, by increasing the tipmaximum thickness of the first blades only, the first blades may be mademore robust against super-cooled ice and may protect the second blades,which in turn may be designed with a tip maximum thickness to optimiseaerodynamic performance and not super-cooled ice impact protection.

The tip maximum thickness of the first blades may greater than 2.7 mm,for example greater than 3.0 mm, or greater than 3.5 mm, or greater than4.0 mm.

The tip maximum thickness of the first blades may be less than 5 mm, forexample less than 4.5 mm, or less than 4.0 mm.

The tip maximum thickness of the second blades may be between 1.2 mm and2.25 mm.

The ratio of the tip maximum thickness of the first blades to themaximum leading edge radius of the first blades may be less than 6.5.

The ratio of the tip maximum thickness of the first blades to themaximum leading edge radius of the first blades may be greater than 2,for example greater than 2.5, or greater than 3, or greater than 3.5, orgreater than 4.

The compressor may comprise two or more stages. For example, thecompressor may comprise three or four stages. The compressor maycomprise less than twelve stages, for example less than eleven, or tenstages.

For example, the compressor may comprise 2 to 8 stages.

The compressor may be an intermediate pressure compressor and the gasturbine engine may further comprise a high pressure compressordownstream of the intermediate pressure compressor.

The turbine may be an intermediate pressure turbine and the gas turbineengine may further comprise a high pressure turbine upstream of theintermediate pressure compressor.

The shaft may be a first shaft and the gas turbine engine may furthercomprise a second shaft coupling the high pressure turbine to the highpressure compressor.

According to a second aspect, there is provided a gas turbine enginecomprising:

-   -   a fan mounted to rotate about a main longitudinal axis,    -   an engine core, comprising in axial flow series a compressor, a        combustor, and a turbine coupled to the compressor through a        shaft,    -   a reduction gearbox that receives an input from the shaft and        outputs drive to the fan so as to drive the fan at a lower        rotational speed than the shaft,        wherein the compressor comprises a first stage at an inlet and a        second stage, downstream of the first stage, comprising        respectively a first rotor with a row of first blades and a        second rotor with a row of second blades, the first and second        blades comprising respective leading edges, trailing edges and        tips,        wherein the ratio of a tip maximum thickness of the second        blades to a tip maximum thickness of the first blades is less        than 0.45.

The ratio of the tip maximum thickness of the second blades to the tipmaximum thickness of the first blades may be less than 0.40, or lessthan 0.35, or less than 0.3. For example, the ratio of the tip maximumthickness of the second blades to the tip maximum thickness of the firstblades may be greater than 0.20, or greater than 0.25, or greater than0.30. For example, the ratio of the tip maximum thickness of the secondblades to the tip maximum thickness of the first blades may be comprisedbetween 0.20 and 0.45, or between 0.25 and 0.45, or between 0.30 and0.45, or between 0.35 and 0.45, or between 0.40 and 0.45, or between0.20 and 0.40, or between 0.25 and 0.40, or between 0.30 and 0.40..

The present disclosure provides alternative solutions to the problem ofice accretion and ingestion, based on increasing the thickness of thefirst rotor blades only, with a limited effect on efficiency, ratherthan increasing the thickness of the blades of all rotors to reduceblade deflection, which would be particularly detrimental to efficiency.

In other words, the present disclosure provides for a solution withhigher overall efficiency, which is more tolerant to core bird impact,less susceptible to damage, and that does not require complex and/orheavy systems to reduce ice accretion.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core may furthercomprise a second turbine, a second compressor, and a second core shaftconnecting the second turbine to the second compressor. The secondturbine, second compressor, and second core shaft may be arranged torotate at a higher rotational speed than the first core shaft.

In such an arrangement, the second compressor may be positioned axiallydownstream of the first compressor. The second compressor may bearranged to receive (for example directly receive, for example via agenerally annular duct) flow from the first compressor.

The gearbox may be arranged to be driven by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example the first core shaft in the example above). For example,the gearbox may be arranged to be driven only by the core shaft that isconfigured to rotate (for example in use) at the lowest rotational speed(for example only be the first core shaft, and not the second coreshaft, in the example above). Alternatively, the gearbox may be arrangedto be driven by any one or more shafts, for example the first and/orsecond shafts in the example above.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan andcompressor(s). For example, the combustor may be directly downstream of(for example at the exit of) the second compressor, where a secondcompressor is provided. By way of further example, the flow at the exitto the combustor may be provided to the inlet of the second turbine,where a second turbine is provided. The combustor may be providedupstream of the turbine(s).

The or each compressor (for example the first compressor and secondcompressor as described above) may comprise any number of stages, forexample multiple stages. Each stage may comprise a row of rotor bladesand a row of stator vanes, which may be variable stator vanes (in thattheir angle of incidence may be variable). The row of rotor blades andthe row of stator vanes may be axially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38, 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm (around 90.5 inches), 235 cm (around92.5 inches), 240 cm, 250 cm (around 100 inches), 260 cm, 270 cm (around105 inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300cm (around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm(around 130 inches), 340 cm (around 135 inches), 350 cm, 360 cm (around140 inches), 370 cm (around 145 inches), 380 (around 150 inches), cm 390cm (around 155 inches), 400 cm, 410 cm (around 160 inches) or 420 cm(around 165 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds), for example in the range of from240 cm to 280 cm or 330 cm to 380 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.30, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allvalues being dimensionless). The fan tip loading may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds), for example in the range offrom 0.28 to 0.31 or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 12 to 16, 13 to 15, or 13 to 14. The bypassduct may be substantially annular. The bypass duct may be radiallyoutside the core engine. The radially outer surface of the bypass ductmay be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹ s, 105 Nkg⁻¹ s, 100 Nkg⁻¹ s, 95 Nkg⁻¹ s, 90 Nkg⁻¹ s, 85 Nkg⁻¹ s or80 Nkg⁻¹ s. The specific thrust may be in an inclusive range bounded byany two of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹ s to100 Nkg⁻¹ s, or 85 Nkg⁻¹ s to 95 Nkg⁻¹ s. Such engines may beparticularly efficient in comparison with conventional gas turbineengines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of, the gas turbine engine that providesa thrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

BRIEF DESCRIPTION OF THE DRAWINGS

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 4 is an enlarged schematic view of a part of a compressor of thegas turbine engine;

FIG. 5 is a partial schematic view, in cross-section, of a first rotorblade illustrating a difference between a minimum and a maximum leadingedge radius;

FIG. 6 is a partial schematic view, in cross-section, showing thedifference between the maximum leading edge radius of a first rotorblade and a maximum leading edge radius of a second rotor blade.

FIGS. 7a and 7b are top views of tips of a first rotor blade and asecond rotor blade, respectively.

DETAILED DESCRIPTION OF THE DISCLOSURE

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area. Whilst the described example relates to aturbofan engine, the disclosure may apply, for example, to any type ofgas turbine engine, such as an open rotor (in which the fan stage is notsurrounded by a nacelle) or turboprop engine, for example. In somearrangements, the gas turbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

FIG. 4 illustrates a forward part of the low pressure compressor 14 infurther detail.

The low pressure compressor 14 comprises a first stage ST1 with a firstrotor R1 and a first stator S1, and a second stage ST2 with a secondrotor R2 and a second stator S2. The low pressure compressor 14 maycomprise other stages, not illustrated.

Each rotor (R1, R2) and stator (S1, S2) comprises a plurality of blades(B1, B2) and vanes (V1, V2), respectively.

In detail, the first rotor R1 and the second rotor R2 comprise a row offirst blades Bland second blades B2, respectively; whereas the firststator S1 and the second stator S2 comprise a row of first vanes V1 andsecond vanes V2, respectively.

The first blades B1 may have a span comprised between 140 mm and 220 mm,and a true chord comprised between 80 mm and 160 mm.

The second blades B2 may have a span comprised between 120 mm and 180mm, and a true chord comprised between 55 mm and 85 mm.

Each blade B1, B2 and vane V1, V2 comprise a root (not illustrated) andan aerofoil portion with a leading edge, a trailing edge and a tip.

The first blade B1 has a leading edge 50, a trailing edge 52, and a tip54.

The leading edge 50 has a leading edge radius variable along the spanbetween a minimum leading edge radius R1min and a maximum leading edgeradius R1max. In FIG. 5, there are schematically illustrated the minimumleading edge radius R1min and the maximum leading edge radius R1max byoverlapping corresponding cross-sections 56, 58 of the first blades B1taken along lines A-A and B-B of FIG. 4, respectively, in such a waythat the leading edge 50 at those section coincide.

The section 56 containing the minimum leading edge radius R1min,illustrated in dashed line in FIG. 5, may be at a span height between20% and 30%, where 0% corresponds to the root and 100% corresponds tothe tip.

The section 58 containing the maximum leading edge radius R1max may beat a span height between 70% and 100%.

The minimum leading edge radius R1min may be greater than 0.20 mm, forexample equal to 0.25 mm.

The maximum leading edge radius R1max may be greater than 0.4 mm, forexample equal to 0.7 mm.

The ratio of the maximum leading edge radius R1max of the first blade B1to the minimum leading edge radius of the first blade B1 may be greaterthan 2.2, for example equal to 2.8.

The second blade B2 has a leading edge 70, a trailing edge 72, and a tip74. Analogously to the first blade B1, the leading edge 70 has a leadingedge radius variable along the span between a minimum leading edgeradius (not illustrated) and a maximum leading edge radius R2max, whichis smaller than the maximum leading edge radius R1max of the first bladeB1. The maximum leading edge radius R2max of the second blade B2 may beat a cross section 78, corresponding to a span height between 85% and100%.

In FIG. 6 there are illustrated the maximum leading edge radius R1max ofthe first blade B1 and the maximum leading edge radius R2max of thesecond blade B2 by superimposing section 58 and section 78 taken alonglines C-C of FIG. 4, respectively, in such a way that the leading edge50 of the first blade B1 and the leading edge 70 of the second blade B2coincide. Section 78 containing the maximum leading edge radius R2max ofthe second blade B2 is illustrated in dashed line in FIG. 6.

The maximum leading edge radius R2max of the second blade B2 may becomprised between 0.1 mm and 0.2 mm, for example equal to 0.16 mm.

According to the disclosure, the ratio of the maximum leading edgeradius R1max of the first blade B1 to the maximum leading edge radiusR2max of the second blade B2 may be greater than 2.8. In an example, themaximum leading edge radius R1max of the first blade B1 may be equal to0.7 mm and the maximum leading edge radius R2max of the second blade B2may be equal to 0.16, such that the ratio of the maximum leading edgeradius R1max of the first blade B1 to the maximum leading edge radiusR2max of the second blade B2 may be equal to about 4.4.

In FIGS. 7a and 7b there are illustrated top views of the tips 54, 74 offirst blade B1 and the second blade B2, respectively.

The tip 54 of the first blade B1 features a maximum thickness T1max thatmay be greater than 2.7 mm, for example equal to 4.3 mm. The maximumthickness T1max may be arranged at a chordwise position between 48% and54%, for example between 50% and 52%, or about 51%, where 0% correspondsto the leading edge 50 and 100% corresponds to the trailing edge 52.

The tip 74 of the second blade B2 features a maximum thickness T2maxthat may be greater than 1.2 mm and less than 2.25 mm, for example equalto 1.7 mm. The maximum thickness T2max may be arranged at a chordwiseposition between 42% and 62%, for example between 48% and 54%, orbetween 50% and 52%, or about 51%, where 0% correspond to the leadingedge 70 and 100% corresponds to the trailing edge 72.

In an example, the tip maximum thickness T2max of the second blade B2 isequal to 1.7 mm, and the tip maximum thickness T1max of the first bladeB1 is equal to 4.3 mm, such that their ratio is equal to about 0.40.

In another example, the tip maximum thickness T2max of the second bladeB2 is equal to 1.3 mm, and the tip maximum thickness T1max of the firstblade B1 is equal to 3.0 mm, such that their ratio is equal to about0.43.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

We claim:
 1. A gas turbine engine comprising: a fan mounted to rotateabout a main longitudinal axis, an engine core, comprising in axial flowseries a compressor, a combustor, and a turbine coupled to thecompressor through a shaft, a reduction gearbox that receives an inputfrom the shaft and outputs drive to the fan so as to drive the fan at alower rotational speed than the shaft, wherein the compressor comprisesa first stage at an inlet comprising a first rotor with a row of firstblades, the first blades comprising respective leading edges, trailingedges and tips, wherein a ratio of a maximum leading edge radius of thefirst blades to a minimum leading edge radius of the first blades isless than
 7. 2. The gas turbine engine according to claim 1, wherein theratio of the maximum leading edge radius of the first blades to theminimum leading edge radius of the first blades is greater than 2.2. 3.The gas turbine engine according to claim 1, wherein the maximum leadingedge radius of the first blades is greater than 0.4 mm.
 4. The gasturbine engine according to claim 1, wherein the maximum leading edgeradius of the first blades is less than 0.9 mm.
 5. The gas turbineengine according to claim 1, wherein the minimum leading edge radius ofthe first blades is greater than 0.15 mm.
 6. The gas turbine engineaccording to claim 1, wherein the minimum leading edge radius of thefirst blades is less than 0.6 mm.
 7. The gas turbine engine according toclaim 1, wherein the compressor comprises 2 to 8 stages.
 8. The gasturbine engine according to claim 1, wherein the compressor is anintermediate pressure compressor, the gas turbine engine furthercomprising a high pressure compressor downstream of the intermediatepressure compressor; the turbine is an intermediate pressure turbine,the gas turbine engine further comprising a high pressure turbineupstream of the intermediate pressure compressor; and the shaft is afirst shaft, the gas turbine engine further comprising a second shaftcoupling the high pressure turbine to the high pressure compressor.
 9. Agas turbine engine comprising: a fan mounted to rotate about a mainlongitudinal axis, an engine core, comprising in axial flow series acompressor, a combustor, and a turbine coupled to the compressor througha shaft, a reduction gearbox that receives an input from the shaft andoutputs drive to the fan so as to drive the fan at a lower rotationalspeed than the shaft, wherein the compressor comprises a first stage atan inlet and a second stage, downstream of the first stage, comprisingrespectively a first rotor with a row of first blades and a second rotorwith a row of second blades, the first and second blades comprisingrespective leading edges, trailing edges and tips, wherein a ratio of atip maximum thickness of the first blades to a maximum leading edgeradius of the first blades is less than 6.5.
 10. The gas turbine engineaccording to claim 9, wherein the ratio of the tip maximum thickness ofthe first blades to the maximum leading edge radius of the first bladesis greater than
 2. 11. The gas turbine engine according to claim 9,wherein the maximum leading edge radius of the first blades is locatedin an area between 70% and 100% of a span height, preferably between 80%and 100%, where 0% corresponds to a root of the first blades and 100%corresponds to the tip.
 12. The gas turbine engine according to claim 9,wherein a ratio of the maximum leading edge radius of the first bladesto a minimum leading edge radius of the first blades is comprisedbetween 2.2 and
 7. 13. The gas turbine engine according to claim 12,wherein the minimum leading edge radius of the first blades is locatedin an area less than 50% of a span height, for example less than 40%, orless than 30%, or between 20% and 30%, where 0% corresponds to a root ofthe first blades and 100% corresponds to the tip.
 14. The gas turbineengine according to claim 9, wherein the tip maximum thickness of thefirst blades is greater than 2.7 mm.
 15. The gas turbine engineaccording to claim 9, wherein the tip maximum thickness of the firstblades is less than 5 mm.
 16. The gas turbine engine according to claim9, wherein the compressor comprises 2 to 8 stages.
 17. The gas turbineengine according to claim 9, wherein the compressor is an intermediatepressure compressor, the gas turbine engine further comprising a highpressure compressor downstream of the intermediate pressure compressor;the turbine is an intermediate pressure turbine, the gas turbine enginefurther comprising a high pressure turbine upstream of the intermediatepressure compressor; and the shaft is a first shaft, the gas turbineengine further comprising a second shaft coupling the high pressureturbine to the high pressure compressor.
 18. The gas turbine engineaccording to claim 9, wherein a ratio of a tip maximum thickness of thesecond blades to the tip maximum thickness of the first blades is lessthan 0.45.
 19. The gas turbine engine according to claim 18, wherein theratio of the tip maximum thickness of the second blades to the tipmaximum thickness of the first blades is greater than 0.10.
 20. The gasturbine engine according to claim 18, wherein the tip maximum thicknessof the second blades is between 1.2 mm and 2.25 mm.